1. Field of the Invention
The invention relates to a rotating assembly for a turbine engine, such as in particular an aircraft jet engine, and a turbine engine including such an assembly.
2. Description of the Related Art
The pre-existing state of the art, which the invention has provided developments and non-negligible advantages to, is described hereunder, while referring to FIGS. 1 and 2. Although the invention can be adapted to the various stages of a turbine engine, it shall be illustrated when it is provided on one stage of a high-pressure turbine. As a matter of fact, the invention is most interesting in such an environment.
According to the known state of the art, a high-pressure turbine thus comprises a disc 10, and blades 12 radially extending outwards from the disc and roots of which, the inner ends of which are referenced 14, are axially engaged in slots, the bottoms of which slots are referenced 16, of the outer periphery of the disc, and radially held by teeth of the disc, the outer ends of which are referenced 18, alternating with said slots. So-called slot cavities 20a, 20b are formed by spaces radially located between the walls of the blade roots 14 and the walls of the slots 16, and which axially extend in the downstream direction of the slots. When rotating, and as a result of the centrifugal effects, the blades are radially pushed outwards, against the side flanks of the teeth 18. Such flanks are also called disc teeth faces. On the high-pressure turbine stage shown, the roots 14 of the blades 12 have two radial stages, an outer one 14a and an inner one 14b, with each of such root stages being held by a pair of faces from two adjacent teeth 18 of the disc also consisting of two stages, an outer one 18a and an inner one 18b. The roots 14 and the teeth 18 thus each are shaped as two radially stacked bulbs. Teeth 18 and roots 14 are also called “fir-tree” elements. In such a configuration outer slot cavities 20a are formed between the outer stages of the roots 14a and the teeth 18a, and inner slot cavities 20b are formed between the inner stages of the roots 14b and the teeth 18b. 
The expression “axially” is to be considered in relation to the longitudinal axis 50 of the turbo machine. As a consequence, “substantially axially” means a direction substantially parallel to said axis 50, to the extent of more or less 10°-15°, globally along which gases (F) flow and around which the rotating assembly rotates.
The blades 12 also comprise inner platforms 22 which extend laterally and which are circumferentially arranged end to end so as to define, together, the inner cylindrical or tapered limit of the hot gas flow circulating in the turbine. The part of the blade 12 located inside relative to the jet, i.e. between the inner platform 22 and the root 14, is called a stilt 24. With such positioning, spaces are formed between two adjacent stilts 24 in the circumferential direction, and between the platforms 22 and the teeth 18 in the radial direction, and form so-called inter-stilts or inter-blade cavities 26. The platforms 22 each have substantially radial walls 22a, 22b, in addition to a cylindrical or tapered wall 22c, which extend therefrom inwards from the respective upstream and downstream ends thereof, so as to partially cover and isolate the upstream and the downstream of the inter-blade cavities 26. One opening however remains between the teeth 18 and the walls 22a, 22b of the platforms 22, so that a flow can axially circulate through the inter-blade cavities.
An upstream annular shroud 28 is provided upstream of the disc 10; Such shroud has an annular hook 30a engaged with an annular hook 30b of the upstream face of the disc 10, and the inner end of the shroud is further bolted to an upstream flange of the disc 10 (such bolting is not shown in the figure). The outer end of the shroud 28 is arranged against the upstream faces of the teeth 18 of the disc and the roots 14 of the blades 12, so that the shroud 28 axially holds the blades 12 in the slots 16 of the disc 10. More particularly, the outer end of the shroud 28 comprises an annular lip 32 protruding downwards, which rests against the above-mentioned upstream faces. Resting may not be perfect, because of the clearances between the parts. As resting is radially located between the outer slot cavities 20a and the inner slot cavities 20b, sealing is created between such two series of slots, upstream of the disc.
A downstream annular shroud 34 is provided downstream of the disc 10; Such downstream shroud 34 is held on the downstream face of the disc 10 by annular hook systems 36a at the inner end of the shroud cooperating with annular hooks 36b of the downstream face of the disc 10. Such downstream shroud 34 comprises an outer annular lip 38 protruding in the upstream direction, and located opposite, or even resting against, the downstream ends of the platforms 22, and more particularly the downstream radial walls 22b. Such downstream shroud 34 also comprises an inner annular lip 40 protruding in the upstream direction, and located opposite, or even resting against the downstream faces of the teeth 18 and the roots 14 of the blades, radially between the outer slot cavities 20a and the inner slot cavities 20b. The inner lip 40 makes it possible to create sealing between the outer slot cavities 20a and the inner slot cavities 20b, downstream of the disc. The outer lip 38 makes it possible to create sealing between the jet and the inter-blade cavities 26 downstream of the blades. With such an arrangement, the downstream shroud 34 also aims at axially holding the blades 12 in the slots 16 of the disc 10.
With such an arrangement, it can be seen that a flow can circulate between the inter-blade cavities 26 and the outer slot cavities 20a, from the upstream or downstream of the disc 10, whereas the inner slot cavities 20b are totally isolated from the other cavities 26, 20a, by the lip 32 of the upstream shroud 28 and the inner lip 40 of the downstream shroud 34.
As mentioned above, the bladed disc discussed here is mounted in the high-pressure turbine of a turbine engine. This is the reason why it can be seen in the figure that it is positioned just downstream of a combustion chamber 42 and of a high-pressure distributor 44, conventionally known from the prior art.
In order to increase the performances of the turbine engine, and to avoid the heating of the disc 10 and the upstream shroud 28 by the flow of hot gases from the upstream combustion chamber and flowing through the jet, it is important to limit as much as possible the flowing of hot gases from the combustion chamber 42 inwards, between the high-pressure distributor 44 and the bladed disc. As a matter of fact, such two stages are axially separated by a certain distance, which forms an annular-shaped discontinuity 46 at the inner limit of the gas jet. Such gases could theoretically flow inwards through such discontinuity 46 and damage the turbine engine. For this purpose, pressurized cold air is taken-off upstream of the combustion chamber in a low-pressure or high-pressure compressor stage, and is transferred to the annular space 46 upstream of the disc 10 and downstream of the high-pressure distributor by a circuit 51 inside the jet. More precisely, a portion of the pressurized cold air (arrow 1) is transferred upstream of the upstream shroud 28 and the other portion (not shown) between the upstream shroud 28 and the disc 10.
The portion of pressurized cold air (arrow 1) which is transferred upstream of the upstream shroud 28 thus flows outwards, along the shroud 28, towards the annular discontinuity 46 of the jet, thus cooling the shroud 28 and the upstream faces of the disc teeth 18, while the pressure and rate thereof prevent the jet gases from flowing inwards, through same discontinuity 46 (arrow 3). The same portion of pressurized cold air circulates in the outer slot cavities 20a (arrow 2) to better cool the outer periphery of the disc 10, on the whole axial extent thereof.
The portion of the pressurized cold air which is transferred downstream of the upstream shroud 28, between same shroud 28 and the disc 10, circulates in the inner slot cavities 20b and supplies a series of conduits (not shown) formed inside the blades 12, and more particularly opening on the trailing edges, leading edges, suction sides and pressure sides thereof. Such conduits cool down the blades 12, which enables these to resist the hot gases from the combustion chamber 42.
In FIG. 1, references 55, 57 indicate conduits which may usually form transferring means for such cooling air, up to the above-mentioned areas, and more specifically the slot cavities.
A series of studies and tests conducted on such architecture made it possible to demonstrate that the cooling air which circulates in the downstream direction through the outer slot cavities 20a then goes up to the outside at the periphery of the disc, along the downstream shroud 34 (arrow 5) and in fact recirculates in the upstream direction through the inter-blade cavities (arrow 4), to be ejected close to the jet discontinuity 46 (arrow 6). However, when the cooling air circulates along such circuit, its calorie content increases and it warms up, when in contact with the platforms, for instance, and transmits such heat to the cooling air having directly flown to the discontinuity 46. In order to keep an acceptable temperature at the periphery of the disc 10 and at the upstream shroud 28, a rather high cooling air rate had to be supplied so far, so as to compensate for such useless heat acquisition through the outer slot cavities 20a and the inter-blades cavities 26.